Gas turbine engine augmenter liner coolant flow control system

ABSTRACT

A stabilizing and support system for an augmenter cooling liner of a gas turbine engine is shown to include flange means which divide a cooling plenum formed between the cooling liner and an exhaust duct casing into a plurality of individual coolant chambers such that the pressure differential across the cooling liner can be closely controlled. The stabilizing and support system includes a plurality of stabilizers adapted to mount the cooling liner to the exhaust duct casing in such a manner as to define the cooling plenum therebetween. The stabilizers are captured on their outer ends by a stabilizer guide which permits relative thermal expansion to occur between the cooling liner and the exhaust duct casing, and each of the stabilizer guides is connected to a positioning band which is adapted to be mounted to the inside of the exhaust duct casing. In the preferred embodiment, the positioning band is provided with an integrally formed flange which defines a restricted inlet to each of the cooling chambers.

I United States Patent 1 1 1111 3,866,417

velegol 14 1 Feb. 18, 1975 GAS TURBINE ENGINE AUGMENTER [57] ABSTRACTLINER COOLANT FLOW CONTROL SYSTEM A stabilizing and support system foran augmenter [75] lnventor: David A. Velegol, Colliers, W, Va. coolingliner of a gas turbine engine is shown to inelude flan e means whichdivide a coolin lenum [73] Abblgnee g-enetrahE-lecmtc Company formed beiween the cooling liner and an exh au st duct incinnati, Ohio casing intoa plurality of individual coolant chambers (22] Filed: Feb. 9, 1973 suchthat the pressure differential across the cooling liner can be closelycontrolled. The stabilizing and l Appl' 3314,78 support system includesa plurality of stabilizers adapted to mount the cooling liner to theexhaust duct [52] U.S. Cl 60/261, 60/39.32, 60/3966 casing in such amanner as to define the cooling [51 1 Int. Cl F02c 7/20, F03k 3/10plenum therebetween. The stabilizers are captured on {58] Field ofSearch 60/261, 39.65, 39.66, 39.69, their outer ends by a stabilizerguide which permits 60/3972 R, 270 R, 39.32, 39.31 relative thermalexpansion to occur between the cooling liner and the exhaust ductcasing, and each of the [56] References Cited stabilizer guides isconnected to a positioning band UNITED STATES PATENTS which is adaptedto be mounted to the inside of the 2,510,645 6/1950 McMahan 60/3932exhaust duct Casing' t Prelerred.embdimem the 2,974,486 3/1961 Edwards60/261 x Positioning band is Pmvded "tegrally formed 3,031,844 5/1962TOmOlOniUS 60/3931 x flange which defines a restricted ihlel 10 each ofthe 3,540,216 11/1970 Quillevere et a1 60/3972 R cooling chambers.3,570,241 3/1971 Alexander 3,712,062 4/1968 Nash 60/3965 X PrimaryExaminerC. J. Husar Assistant Examiner-Robert E. Garrett Attorney,Agent, or Firm-Derek P. Lawrence; Lee H. Sachs 8 Claims, 3 DrawingFigures PATENTED FEB I 8 1975 add-- GAS TURBINE ENGINE AUGMENTER LINERCOOLANT FLOW CONTROL SYSTEM BACKGROUND OF THE INVENTION This inventionrelates generally to augmented gas turbine engines and, moreparticularly, to means for controlling and regulating augmenter linercoolant pressure.

The invention herein described was made in the course of or under acontract, or a subcontract thereunder, with the United States Departmentof the Air Force.

Gas turbine engines generally comprise a compressor for compressing airflowing through the engine, a combustion system in which high energyfuel is mixed with the compressed air and ignited to form a high energygas stream, and a turbine which includes a rotor portion operativelyconnected to the compressor to drive the same. Many modern-day gasturbine engines are of the turbofan type in which a second or lowpressure compressor is mounted forwardly of the high pressure compressorand is driven by a second turbine mounted downstream of the firstturbine. The low pressure compressor or fan presents an additional stageof compression and, in addition, is normally of a larger diameter thanthe high pressure compressor. The turbofan engine is therefore capableof flowing a much larger mass of air, thereby greatly increasing thethrust output of the engine.

An additional known method of increasing the thrust output of the engineis to provide the engine with an augmenter. In such an engine,additional fuel is injected into an exhaust duct formed downstream ofthe second turbine and is ignited to provide an additional high energygas stream, which is ejected through an exhaust nozzle to provide highenergy thrust output from the engine. In the case of turbofan engines,it is also known to mix the fan airflow with core engine airflow priorto supplying the mixed flow with additional fuel by means of theaugmenter.

The augmenter system is normally located within the exhaust duct of theengine and, in most cases, some means must be provided for protectingthe exhaust duct from the extremely high temperatures associated withthe gas flow within the augmenter. One common means for providing thisprotection is to position a cooling liner within the exhaust duct and topass cooling air between the liner and the exhaust duct. Openings orslots are positioned within the cooling liner such that the cooling airmay flow through these openings and form a film of coolant on the insideof the cooling liner.

A basic problem confronting the designer of such an augmenter linerconsists of the regulation of the coolant flow and the minimization ofliner pressure loading. In order to use the minimum amount of coolingair for maximum engine performance, a simple and reliable means isneeded to regulate the coolant flow through the openings or slots in thecooling liner. The problem is complicated by the fact that the pressurein the plenum which surrounds the cooling liner is essentially constantalong the entire axial length of the liner while the static pressure ofthe combustion gas inside the liner decreases axially due to theacceleration of the gas as its temperature increases due to operation ofthe augmenter. This condition not only results in significant pressuredifferentials across the liner but also in a varying pressuredifferential from the upstream to the downstream end of such liner.

In order to offset the increasing pressure differential, designers have,in the past, attempted to maintain an essentially constant coolant flowby varying the size of the openings and slots within the liner. In orderto offset the higher pressure loading at the aft end of the liner,various means such as reinforcing rings or additional mounting pointsand hangers have been suggested. Obviously, the additional complexityinvolved in varying the size of the coolant openings can easily increasethe manufacturing costs of such a liner. In addition, the requirementsfor heavy stabilizing rings and- /or additional mounting hardware mayunduly increase the weight of the overall system.

SUMMARY OF THE INVENTION It is an object of the present invention,therefore, to overcome the above-mentioned problems and to provide meansfor regulating the coolant flow to a plenum surrounding a gas turbineengine augmenter cooling liner such that a relatively constant pressuredifferential can be achieved along the entire axial length of the liner.It is an additional object of this invention to provide such a coolantregulating means which yields relatively uniform coolant flow throughopenings provided in the liner without the necessity of varying the sizeof such openings along the axial length of the liner.

Briefly stated, the above and similarly related objects are attained inthe present instance by providing a cooling liner stabilization andmounting system which includes means which divide the coolant plenumsurrounding the liner into a number of individual chambers. The flowinto each of these chambers is con trolled by means of a flangeassociated with the mount ing system which acts as a restriction in theflow path and also divides the plenum into the previouslymentionedindividual chambers. By varying the size of the flange, the pressurewithin the individual chambers can be readily controlled and thepressure differential across the liner can thereby readily be regulated.

DESCRIPTION OF THE DRAWING While the specification concludes with aseries of claims which particularly point out and distinctly claim thesubject matter which Applicant regards as his invention, a clearunderstanding of the invention will be obtained from the followingdetailed description, which is given in connection with the accompanyingdrawing,

in which:

FIG. 1 is a schematic, axial, cross-sectional view of a gas turbineengine incorporting the present inventive regulating means;

FIG. 2 is an enlarged, partial view of the inventive cooling linerstabilization and pressure regulation means of FIG. 1; and

FIG. 3 is a graphical plot showing the pressure differential along theaxial length of a cooling liner incorporating the present inventivemeans.

DESCRIPTION OF A PREFERRED EMBODIMENT Referring to the drawing whereinlike numerals correspond to like elements throughout, attention isdirected initially to FIG. 1 wherein a gas turbine engine 10 of themixed flow turbofan type is shown to include a core engine 12 whichincludes a fan turbine 14 for driving a plurality of fan blades 15mounted on a shaft 16. The fan blades are located within an inlet 17formed by an outer or fan casing 18 which surrounds the entire gasturbine engine 10. The fan casing 18 operates with a core engine casing20 to define parallel flow paths 22 and 23.

Air entering the flow path 23 is compressed by means of a compressor 24and is mixed with fuel in combustor 26. Fuel is delivered to thecombustor 26 by means of a plurality of fuel injection points 27 fromfuel tubes 28 which extend through the flow path 22. The resultant highenergy gas stream exits the combustor 26 and drives a turbine 30 which,in turn, drives the compressor 24 by means of a shaft 31.

As further shown in FIG. 1, air flowing through the outer or fan flowpath 22 and air exiting the core engine 12 flow through a mixer 32,which operates to mix the two separate flow paths. The mixed flow pathis then acted upon by an augmenter 34, which consists of a plurality offuel injectors 36. The resultant fuel/air mixture in the augmenter 34 isignited by means of a suitable igniter (not shown), flows through anexhaust duct 40, and thereafter provides an additional propulsive forceby exiting through an exhaust nozzle 42.

The exhaust duct is located at the downstream end of the fan casing 18and is shown in FIG. 1 to include an exhaust duct casing 44 and acooling air liner which is generally designated by the numeral 46. Thecooling liner 46 is spaced radially inwardly from the exhaust ductcasing 44 and defines an annular coolant flow path 48 having an inlet 50formed by a forward lip 52 at the upstream end of the cooling liner 46.

As is well known in the art, the cooling liner 46 includes a pluralityof openings or slots 54 adapted to deliver cooling air from thepassageway 48 to the inside of the liner 46. The coolant flowing throughthe open ings 54 provides a film of cool air on the inside of the liner46 thereby protecting both the liner 46 and the surrounding exhaust ductcasing member 44 from the high temperatures associated with theoperation of the augmenter 34.

The operation of the engine 10 is well known and will be discussed onlybriefly. Air flows through the inlet 17 and is acted upon by the fanblades 15. A first portion of this pressurized air flows through the fanflow path 22, while a second portion flows through the core engine flowpath 23 and is acted upon by the compressor 24. A high energy gas streamis generated by the combustor 26 and drives the high pressure turbine 30and low pressure turbine 14, which, in turn, drive the core enginecompressor 24 and the fan 15. Air exiting the low pressure turbine 14and air flowing through the fan flow path 22 are mixed within the mixer32 and the mixed flow is delivered to the region of the augmenter 34. Aresultant fuel/air mixture generated by the augmenter 34 is ignited toprovide an additional propulsive force by exiting through the exhaustnozzle 42.

A portion of the air flowing through the fan flow path 22 flows throughthe inlet 50 and, thus, through the coolant passageway 48. This coolingair thereafter flows through the openings 54 and forms a film on theinside of the cooling liner 46 thereby protecting the liner 46 and thesurrounding exhaust duct casing 44 from the high gas temperaturesassociated with operation of the augmenter 34.

The gas turbine engine 10 described above is typical of many present-dayaugmented turbofan engines and has been described solely to place thepresent invention in proper perspective. As will become clear to thoseskilled in the art, the present invention will be applicable to othertypes of gas turbine engines and, therefore, the engine 10 is merelymeant to be illustrative.

Referring now to FIGS. 2 and 3, the gas turbine engine augmenter coolingliner 46 and its associated pressure regulating system is shown ingreater detail. As shown in FIG. 2, the cooling liner 46 is mounted tothe exhaust duct 44 by means of a plurality of stabilizer assemblies 56,the details of which are shown and claimed in application Ser. No.328,769, filed in the name of D. O. Nash et al. and assigned to the sameassignee as the present invention now US. Pat. No. 3,826,088. Asdescribed in the Nash et al. application, the stabilizer assembly 56includes a circumferential row of stabilizers 58, each of which arecaptured within a stabilizer guide 60. The stabilizer guides 60 are, inturn, connected to the inner surface of a positioning band 62, which isadapted to be connected to the inner wall surface of the exhaust duct44.

The positioning band 62 includes a plurality of holes 64 which arealigned with threaded openings associated with capture nuts 68 mountedon each of the stabilizer guides 60. A plurality of bolts 70 fit throughopenings 72 formed within the exhaust duct casing 44 and are threadablyreceived within the nuts 68, thereby firmly attaching the positioningband 62 and thus the cooling liner 46 to the exhaust duct casing 44.

As more fully described in the Nash et al application, the stabilizers58 are adapted to define radial height 71 of the annular coolantpassageway 48. As discussed briefly above, tests with previousaugmenters have shown that the pressure of the coolant within thepassageway 48 is relatively constant along the entire axial length ofthe liner 46. The static pressure of the combustion gas within theexhaust duct, i.e. inside the liner 46, decreases axially due to theacceleration of the gas as its temperature increases due to operation ofthe augmenter. With the relatively constant pressure within thepassageway 48 and the decreasing pressure within the exhaust duct 40,the pressure differential across the liner 46 increases significantlynear the aft end of the liner 46. This condition is shown graphically asa dotted line curve 73 in FIG. 5 wherein pressure differential acrossthe liner is plotted as a function of axial length of the liner.

Referring still to FIGS. 2 and 3, the present invention includes meansfor regulating the pressure level within the coolant passageway 48 suchthat the pressure differential across the cooling liner can bemaintained relatively constant from the upstream end to the downstreamend of the liner. The pressure-regulating means comprise a flange 76formed at either the upstream or downstream end of the positioning band62. While the flange 76 may be made a separate member, in the embodimentshown in FIG. 2 the flange 76 is formed integrally with the positioningband 62. As shown best in FIG. 3, the flanges 76 associated with each ofthe stabilizer assemblies 56 act to divide the coolant passageway 48into separate annular chambers 78, 80 and 82.

The flange 76 is sized so as to provide a gap 84 between an inner edge86 of the flange 76 and the outer wall of the cooling liner 46. The gap84 acts as a restriction in the flow path for the cooling air and actsto control the pressure of the coolant air within the chambers 78, 80and 82. The gap 84 may be made of various sizes depending upon thedesired pressure within the chambers 78, 80 and 82. In certainapplications, the gap 84 may be the same for each of the stabilizerassemblies 56, while in other applications the gap size may vary frommounting assembly to mounting assembly. In either case, the gap 84 canbe utilized to provide a pressure within the chambers 78, 80 and 82 suchthat the average pressure differential across the liner is a minimum ineach chamber. This minimum is limited by the least value of pressuredifferential which will produce the required radially inward coolingflow. This situation is plotted as solid line curve 84 in FIG. 3 withthe least desired pressure differential lying below the point A. Asshown therein, the pressure differential across the liner 46 may varyslightly from the upstream to the downstream end of each of the chambers78, 80 and 82, but the net effect of the invention system is to providea much lower overall pressure differential from the upstream end of theliner to the downstream end thereof. With such a result, the liner canbe made with uniformly sized slots or openings 54 and the need for heavyreinforcing rings at the downstream end thereof becomes non-existent.

It should be obvious to those skilled in the art that slight variationscould be made in the structural elements shown and described abovewithout departing from the broader inventive concepts disclosed herein.For example, the flange 76 could be positioned at either end of thepositioning band 62. Similarly, the flange 76 could be formed integrallywith, or connected to, the stabilizing guides 60 instead of thepositioning band 62. Likewise, the flange 76 could be connected directlyto either the exhaust duct casing 44 or the cooling air liner 46. Inaddition, the stabilizing assemblies 56 are capable of use incombination with more standard mounting techniques such as spacers 86shown in FIG. 3 near the upstream end of the liner 46. The appendedclaims are intended to cover these and similar modifications inApplicants broader inventive concept.

What is claimed is:

1. In a gas turbine engine of the type including a compressor, aturbine, a combustion system, an augmenter, an exhaust duct surroundingsaid augmenter, and a cooling liner positioned within said duct so as toform a cooling plenum therebetween, at least a portion of which linerextends downstream from said augmenter and is adapted to protect saidexhaust duct from the high temperature gas generated by said augmenter,the

6 improvement comprising:

flange means situated within said cooling plenum and dividing saidcooling plenum into at least two individual chambers, said flange meansoccupying a predetermined radial space in said cooling plenum andproviding a restricted inlet between said chambers such that the averagepressure differentials across the liner in each of said chambers aresubstantially equal.

2. The improved gas turbine engine of claim 1 wherein said linerincludes a plurality of stabilizer assemblies mounted between said linerand said exhaust duct in such a manner as to define said coolant plenumand each of said stabilizer assemblies includes a plurality ofstabilizers, a like number of stabilizer guides, one of which partiallysurrounds each of said stabilizers, and a positioning band surroundingand mounted to each of said stabilizer guides and to said exhaust duct,and said flange means comprises a flange formed integrally with saidpositioning band and extending toward said cooling liner.

3. The improved gas turbine engine of claim 1 wherein said inlet betweensaid chambers is sized to provide a relatively constant average pressuredifferential in each of said chambers along the axial length of saidliner.

4. The improved gas turbine engine of claim 2 wherein said cooling linerincludes at least two of said stabilizer assemblies such that saidplenum is divided into at least three of said chambers.

5. The improved gas turbine engine of claim 4 wherein said cooling linerincludes a plurality of relatively equally sized coolant holes spacedalong at least a portion of the axial length thereof.

6. The improved gas turbine engine recited in claim 1 wherein saidinlets are sized so as to provide a pressure differential across saidcooling liner ranging between 0 and 0.5 A P, where A P is defined aspressure in said chamber-pressure in said duct.

7. The improved gas turbine engine recited in claim 6 wherein saidpressure differential ranges between 0 and 0.5 A P in each of saidchambers.

8. The improved gas turbine engine of claim 4 wherein said restrictedinlets are sized and said flange means are axially spaced from oneanother such that the pressure differential across said cooling liner isreduced.

1. In a gas turbine engine of the type including a compressor, aturbine, a combustion system, an augmenter, an exhaust duct surroundingsaid augmenter, and a cooling liner positioned within said duct so as toform a cooling plenum therebetween, at least a portion of which linerextends downstream from said augmenter and is adapted to protect saidexhaust duct from the high temperature gas generated by said augmenter,the improvement comprising: flange means situated within said coolingplenum and dividing said cooling plenum into at least two individualchambers, said flange means occupying a predetermined radial space insaid cooling plenum and providing a restricted inlet between saidchambers such that the average pressure differentials across the linerin each of said chambers are substantially equal.
 2. The improved gasturbine engine of claim 1 wherein said liner includes a plurality ofstabilizer assemblies mounted between said liner and said exhaust ductin such a manner as to define said coolant plenum and each of saidstabilizer assemblies includes a plurality of stabilizers, a like numberof stabilizer guides, one of which partially surrounds each of saidstabilizers, and a positioning band surrounding and mounted to each ofsaid stabilizer guides and to said exhaust duct, and said flange meanscomprises a flange formed integrally with said positioning band andextending toward said cooling liner.
 3. The improved gas turbine engineof claim 1 wherein said inlet between said chambers is sized to providea relatively constant average pressure differential in each of saidchambers along the axial length of said liner.
 4. The improved gasturbine engine of claim 2 wherein said cooling liner includes at leasttwo of said stabilizer assemblies such that said plenum is divided intoat least three of said chambers.
 5. The improved gas turbine engine ofclaim 4 wherein said cooling liner includes a plurality of relativelyequally sized coolant holes spaced along at least a portion of the axiallength thereof.
 6. The improved gas turbine engine recited in claim 1wherein said inlets are sized so as to provide a pressure differentialacross said cooling liner ranging between 0 and 0.5 Delta P, where DeltaP is defined as pressure in said chamber-pressure in said duct.
 7. Theimproved gas turbine engine recited in claim 6 wherein said pressuredifferential ranges between 0 and 0.5 Delta P in each of said chambers.8. The improved gas turbine engine of claim 4 wherein said restrictedinlets are sized and said flange means are axially spaced from oneAnother such that the pressure differential across said cooling liner isreduced.